The present invention relates generally to gas turbine engines, and, more specifically, to fan and compressor rotor blades therein.
In a typical turbofan gas turbine engine, a multistage axial compressor pressurizes air which is mixed with fuel in a combustor and ignited for generating hot combustion gases which flow downstream through a high pressure turbine which extracts energy therefrom for powering the compressor. A fan is disposed in front of the compressor and is powered by a low pressure turbine behind the high pressure turbine.
The fan and compressor stages each includes a row of circumferentially adjoining rotor blades extending radially outwardly from a supporting rotor disk. Each blade includes an airfoil over which the air being pressurized flows, and a platform at the root of the airfoil which defines the radially inner boundary for the airflow.
The blades are typically removable, and therefore include a suitable dovetail configured to engage a complementary dovetail slot in the perimeter of the rotor disk. The dovetails may either be axial-entry dovetails or circumferential-entry dovetails which engage corresponding axial or circumferential slots formed in the disk perimeter. A typical dovetail includes a neck of minimum cross sectional area extending integrally radially inwardly from the bottom of the blade platform which then diverges outwardly into a pair of opposite dovetail lobes or tangs.
For axial dovetails, the rotor disk includes a plurality of circumferentially spaced apart, axially extending dovetail slots defined circumferentially between corresponding disk posts. The axial slots and disk posts extend the full axial thickness of the disk between its axially forward and aft faces.
For a circumferential dovetail, a single dovetail slot extends circumferentially around the entire perimeter of the disk, and axially between forward and aft continuous posts. The circumferential slot is locally enlarged at one location for allowing the individual circumferential dovetails to be initially inserted radially therein and then repositioned circumferentially along the dovetail slot until the entire slot is filled with a full row of the blades.
In both types of dovetails, the corresponding disk posts include complementary lobes or tangs which cooperate with the dovetail lobes to radially retain the individual blades against centrifugal force during operation. Each dovetail lobe includes a radially outwardly facing outer pressure surface or face which engages a corresponding radially inwardly facing pressure surface or face of the disk posts. As centrifugal load is generated by the blade during rotation, it is carried radially outwardly from the dovetail lobes into the corresponding disk posts at the engaging outer and inner pressure faces thereof, and then radially inwardly through the disk.
Since the dovetail necks have minimum cross sectional area between the blade platforms and the dovetails themselves, maximum centrifugal stress is experienced at the necks which must be limited for ensuring a suitable blade life. A typical compressor blade is designed for an infinite life which requires suitably large dovetails and necks thereat for experiencing centrifugal stress suitably below the strength limits of the blade material.
The rotor disks, in contrast, have a finite limited useful life since they are more highly stressed than the blades which they retain. Since axially extending dovetail slots in a disk perimeter interrupt that perimeter along its circumference, an axial-entry rotor disk reacts the centrifugal loads in a different manner than that of the circumferential-entry rotor disks in which the two corresponding disk posts are full circumferential hoops having a correspondingly high hoop strength.
In one type of turbofan aircraft gas turbine engine which entered commercial service in this country in the previous decade, mid-life experience thereof first uncovered high cycle fatigue cracks in a statistically small, yet significant, number of axial dovetails and the rotor disks therefor. This is undesirable since a crack in a single dovetail uncovered in a periodic maintenance inspection requires replacement of a full set of the blades, or replacement of the rotor disk if the crack is found therein instead.
To improve the high cycle fatigue life of the rotor disk for axial dovetails, an improvement in the axial dovetail slots was patented by the present assignee in U.S. Pat. No. 5,141,401-Juenger et al.
Although that same type of engine includes both axial-entry and circumferential-entry dovetails, cracks in the latter were not observed at mid-life inspections in view of the different and stronger configuration of the circumferential-entry dovetails and slots therefor. However, further continued life of the same engines into the present decade have now uncovered yet again a statistically small occurrence of cracking in circumferential dovetails at very high-life cycles or time.
Analysis has determined that locally high stress occurs at both the radially outer and inner edges of contact between the dovetail and post pressure faces through which blade loads are carried. The stress field near the contact edges is further concentrated by small radii fillets in this region.
Accordingly, it is desired to provide an improved rotor blade and cooperating rotor disk for reducing edge of contact stress.